Stiffened laminate and sandwich composite materials used in military aircraft includes multiple layers that exhibit subsurface damage mechanisms such as disbonds, ply cracking, core crushing, and core cracking, which usually cannot be identified using visual inspections. It is also difficult to inspect composite structures such as blades and fuselage sections using localized methods (e.g., coin tap, ultrasonic, thermographic, impedance) unless these structural components are first removed from the aircraft. Furthermore, these localized methods can detect damage within approximately ⅛ inch of the surface, but do not penetrate into deep composite laminate or sandwich sections.
Laser shearography systems are commercially available for inspecting helicopter rotor blades, but these systems require that a vacuum be achieved prior to testing and that the blades be removed from the rotor for inspection. A number of researchers have applied single-dimensional laser vibrometry for detecting damage in metallic and composite materials based on the analysis of modal frequency responses, vibration curvatures, and strain energies in addition to propagating elastic waves. This prior work has focused on thin laminate composite sections with relatively simple geometries and damage mechanisms located close to the surface of the material. This work in damage detection using laser vibrometry has also relied on reference signatures that are acquired in undamaged specimens. As mentioned above, there is a need for inspection methods other than transmitted x-ray for detecting damage in deep composite sections. There is also a need for damage detection algorithms that do not involve a comparison with reference signatures because there are various composite structures throughout the aircraft that are geometrically complex and exhibit subtle changes in boundary conditions throughout the life of the aircraft.
Some embodiments of the present invention pertain to a vibration-based damage detection technique for composite materials by considering nonlinear changes in the forced frequency response of a fiberglass composite panel due to material damage. By utilizing vibration measurements, the technique addresses the need for wide area damage detection in large composite structures. There is also the potential to implement this inspection method without first removing the structures from the aircraft because a scanning laser vibrometer with a relatively long measurement range is used to acquire the vibration data. One embodiment of the present invention pertains to a reference free, non-linear method for detecting damage using vibration data.